Service vehicle for performing in-space operations on a target spacecraft, servicing system and method for using a service vehicle

ABSTRACT

A service vehicle for performing an in-space operation on a selected target spacecraft incldues a communication module having at least one of a transmission and a receiving characteristic configurable in order to meet at least one of a transmission and a receiving parameter of the selected targeted spacecraft. In addition, a servicing system using the service vehicle and a method for in-space servicing are also disclosed.

The invention relates to a service vehicle for performing in-spaceoperations on a target spacecraft. It furthermore relates to a servicingsystem and to a method for in-space servicing of spacecraft.

Spacecraft in general need to be properly positioned in a predeterminedorbit and be properly oriented in the three-dimensional space withrespect to their service areas in order to fulfill their respectivemission. In other words, they typically are designed to have theirtelecommunication equipment looking to (or pointing to) the servicearea. Various forces such as moon gravity, sun gravity, non-uniformityof gravity potential of earth, solar pressure, and atmosphere in lowaltitudes, and even Venus gravity, plus many other less importantforces, interact with the spacecrafts and tend to change their optimumposition and orientation. These sources alter the orbital elements ofthe respective spacecraft effecting what is called orbit perturbations.To counteract these perturbations, spacecraft are provided withthrusters, which are used either in continuous mode or in pulse mode oroccasionally, from time to time (i.e. every a few days/weeks/months).Said thrusters consume fuel in order to effect the counteracting forces.

Artificial satellites are a particular case of spacecraft as theirmission involves orbiting a specific celestial body in order to be ableto provide their service. Other spacecraft have trajectories that maydiffer for part of their mission from the classical definition ofsatellite orbiting but still have a service area where they have topoint to and accordingly may be negatively influenced by similarperturbations. Usually they become satellites of another celestial bodyor simply float in space at a Lagrange point or elsewhere. The samenature of problems pertains to all type of spacecraft as regards theirorbit and health issues. For reasons of clarity, the followingdescription focuses on a satellite in the proximity of earth and inparticular in a proximity that teleoperation capability is not hinderedby long electromagnetic wave propagation times, although the conceptsmay also be relevant to other kinds of spacecraft.

A spacecraft that can be kept, by means of its thrusters, in a desiredtarget postion and attitude is considered under control or controllable,and a non-controllable spacecraft is out of control with regard to itsposition and attitude. Said controllable spacecraft can be more easilyand safely approached for servicing, and is called “co-operative”, whilea spacecraft that has lost its attitude control is called “nonco-operative”.

Typical spacecraft are designed for a so-called “designed lifetime”. The“designed lifetime” of a spacecraft has a statistical definition.Spacecraft are designed to have an operational lifetime of e.g. 10 yearsat minimum, with an associated probability 98% (based on the statisticallifetime of the selected components). This means that in the term of 10years a portion of 2% of the spacecrafts of the same design and materialand processes would fail and the rest would continue to function. Theaverage lifetime of the materials of a spacecraft is much longer,sometimes 3 times the “designed lifetime”. For example, the voyagerspacecraft still operate after 25 years, while most of thetelecommunication satellites have a designed lifetime of 6 to 15 years.

The spacecraft are designed to carry a predetermined amount of fuel,which is calculated in dependence of what they would need to consumeduring their “designed lifetime”. Consequently, a spacecraft carriesfuel only for the designed lifetime (e.g. 10 years) in order to performall types of maneuvers. At a certain point of time, when fuel reservesfinish, a spacecraft cannot retain even its attitude correct and so itbecomes useless.

When the fuel reserves are very limited, then the spacecraft can nolonger provide the same level of service that it was designed for, oreven provide any useful service at all. In this case the spacecraft isrendered useless and abandoned in space creating an additional problemof potential collision with a future operational spacecraft. It becomesas it is called “space debris”.

Fuel-depletion that renders the spacecraft uncontrollable and thereforeuseless, may happen even earlier than the designed lifetime of thespacecraft for various reasons (e.g simple bad calculation of the fuelbudget, wrong positioning due to error, malfunctioning of the launcher,rare phenomena, accident or otherwise, during the launch phase; wrongpositioning of the spacecraft during the LEOP (Launch and Early OrbitPhase) due to error, malfunction, rare phenomena, accident or otherwise;change in mission; errors, malfunctions, rare phenomena, accidents, orotherwise during the remaining actual lifetime).

In general, any component, unit, subsystem of a spacecraft, such assensors, actuators, processing units, inertial subsystems, powersubsystem, software, communication payload, may fail due to errors inits use, malfunction, rare phenomena or otherwise that may render thespacecraft partially or totally, temporarily or permanentlyuncontrollable and therefore unable to function properly to generate theopportunity revenue, or any revenue. It may even create a significantrisk for other spacecrafts by its status as space debris. In thisdefinition of space debris we will add to the traditionally conceivedone, that regards space debris as passive objects, the characteristic ofpotentially active object that may be even more dangerous than a passivedebris, as an active and unpredictable (accelerating, decelerating)moving object may be.

For both reasons, i.e. lifetime restrictions due to limited fuelresources as well as system failure due to unexpected error, servicingcapabilities for spacecraft with the general goal of artificiallyextending the lifetime of a spacecraft are highly desireable,particularly in view of the typically very high costs involved withreplacing an existing spacecraft by a substitute.

Several inventions have been developed for solving the problem ofproviding servicing capabilities for spacecraft, particularly withrespect to failure on satellites and fuel-depletion (U.S. Pat. No.5,410,731, U.S. Pat. No. 5,813,634, WO 0103310), disclose concepts toinspect the satellites on orbit (U.S. Pat. No. 6,296,205, U.S. Pat. No.6,384,860), disclose concepts to provide service to them on orbit (WO9731822, U.S. Pat. No. 4,896,848, U.S. Pat. No. 4,273,305, U.S. Pat. No.5,299,764, U.S. Pat. No. 4,349,837), or prepare for servicing (U.S. Pat.No. 4,946,596, EP 1 101 699, U.S. Pat. No. 4,657,221). Several othershave developed concepts for tools to perform the service (U.S. Pat. No.4,177,964, WO 0208059) or developed methods for providing new services(EP 1 245 967) for which this invention provides improvements.

Various systems have been studied, wherein the method for performing therendezvous typically is carried out by manual Tele-operation. In someother documents, autonomous rendezvous and docking systems are proposed.

In the case of autonomous docking mechanisms, the designs that have beenproposed involve a robotic arm which demands high dry mass and powerbudgets.

Patent U.S. Pat. No. 5,299,764 discloses a system for carrying outin-space servicing of spacecraft, wherein artificial life robotics areutilized.

Patent U.S. Pat. No. 6,296,205 discloses a concept of inspecting the RFfunctioning of a satellite at proximity and emitting control signals anddiagnostics to the ground.

Patent U.S. Pat. No. 6,384,860 discloses a video telemetry system formonitoring the deployment of an apparatus coupled to a satellite. Thisallows the solar panels to be observed during deployment and even beforesaid panels are deployed, but at very low rate (one frame every 27seconds), said rate not permitting any real teleoperation and any otherservice.

In the cases where teleoperated designs of service vehicles are proposedthese are disadvantaged by the high bandwidth required from the servicevehicles to support the teleoperation. To perform an inspection orrendezvous and docking to a satellite a high bandwidth link needs to beestablished for certain minutes or hours in order to provide sufficientand timely (real time) visual information to the operators and systemson earth to perform the servicing (inspection, rendezvous, docking).Such designs have been proposed resulting in the GSV GeostationaryService Vehicle concept spacecraft.

The disadvantages of this category of prior art are:

-   -   High electric power budgets, in order to cope with the required        high bandwidth transmission for transmitting timely (in real        time) the output of the rendezvous sensors (radar, visual        images) towards the ground stations.    -   High mass budget for the Mission Communication payload,        batteries, solar cells, plus structural overhead and overheads        to the attitude control subsystem (flywheels, thrusters . . . ).    -   High volume as result of the above increased budgets (mass,        structural overhead, protruding antennas, protruding solar        panels).    -   High complexity due to the redundancy required.    -   Higher vulnerability to radiation hazards and debris (larger        profile).    -   Low range of operation as regards delta velocity potential.    -   Large consumption of consumables (fuels, pressurization gas).    -   Low maneuverability due to high volume and mass.    -   Higher risks of client due to higher mass and volume and lower        maneuverability.    -   Larger debris problem at end of its life.

The complexity of service missions to orbiting satellites and the highcost involved (space shuttle cost is 500 M$ per flight) has rendered theidea of servicing ailing satellites as a solution to restore or prolongservice unattractive. As an alternative, putting into orbit universalback-up satellites or specifically designed, individual backupsatellites is considered.

The Geostationary satellites in order to reach their orbit need to usesome kind of launch vehicle of which vehicle either the last part (upperstage) or the apogee kick motor is jettisoned in the space close to thegeostationary ring creating space debris. Said debris constitutes a highhazard potential for future missions. Some recent satellites use aUnified Propulsion System for reaching geosynchronous orbit from theirinjection point and for orbit maintenance. This solution saves one pieceof debris but results to higher mass overheads for the duration of theentire life of the satellite. At the end of life of the satellite thetotality of it becomes space debris.

Up to now, almost no spacecraft has been designed to be refueled or beserviced in space. As one result of this design philosophy, a large partof space debris consists of spent spacecrafts and apogee kick motors andupper stages.

Therefore, it is an object of the present invention to provide aparticularly versatile and flexible service vehicle for performingin-space operations on a target spacecraft. Furthermore, a servicingsystem and a method for in-space servicing of spacecraft shall beprovided.

With respect to the service vehicle, this object is achieved with acommunication module which with respect to its transmission and/orreceiving characteristics is configurable in order to meet givenreceiver and/or transmitter parameters of said selected targetspacecraft.

The services provided by the service vehicle may include any types ofservices, such as refueling, delivering all kinds of material, repair ormaintenance services, or other kinds of suitable activities. Saidservices may collectively be denoted as ACR for Assembly, Convert andRepair. The majority of said ACR services are to be performed by meansof teleoperation assisted by stereoscopic means, illuminating means &tape-tools that assist in fetching/storing tools and fetching/storingspares and fetching/storing disassembled components.

The invention is based upon the concept that for flexible and versatileservicing of a target spacecraft, the service vehicle ought to bedesigned for a particularly low mass, energy and/or fuel budget.However, a significant contribution to both mass and energy/fuelrequirements is the necessity to constantly provide for reliablecommunication between the service vehicle and a ground control station,in particular in view of the comparatively large distances that must beovercome between ground control and service vehicles in expectedservicing missions. In order to significantly lower the onboard powerconsumption on the service vehicle necessary for maintaining a reliablecommunication channel with ground control, the service vehicle isdesigned for emitting signals to and/or receiving signals from groundcontrol by using the target spacecraft to be serviced as a relaystation. In this concept, the energy required from the service vehiclemay be limited to maintain a communication channel with the targetspacecraft, and accordingly the mass required to provide these loweredenergy levels—i.e. accumulator mass—may be kept correspondingly low. Themajor share of the energy necessary to maintain proper communication inthis concept is then delivered by the target spacecraft which as such isdesigned for communicating with ground control anyway. In order torender the target spacecraft useable for this purpose, the servicevehicle is designed to be configurable to establish communicationcontact with the target spacecraft.

Particularly advantageous features of the present invention arespecified in the dependent claims.

Preferably, the configurable communication module comprises atransmitter to emit communication signals to the target spacecraft.Alternatively or additionally, in a preferred embodiment thecommunication module is equipped with an adjustable or configurablereceiving unit, thus allowing the communication module to be set up toreceive input or command signals from various selectable sources. In aparticularly cost- and budget-effective setup, the receiver preferablyis designed adjustable in its working frequency in order to communicatewith a telemetry channel of said selected target spacecraft. In thisconfiguration, the so-called telemetry channel, which in typicalspace-craft designs for safety reasons has a comparatively wide-spreademission characteristics, may be used to relay command or input signalsfrom ground control to the service vehicle. In this way, availablefrequency ranges may be used in a very efficient way for communicationwith the service vehicle.

In a preferred embodiment, the service vehicle is designed withparticular emphasis on the concept to keep the target spacecraft safefrom over-extensive or potentially destructive energy input from theservice vehicle while also providing for a comparatively high range ofdistances to the targed spacecraft over which reliable communication maybe established. In order to achieve these accumulated goals, which withrespect to the output power emitted by the service vehicle contradicteach other, the service vehicle preferrably is designed for variableoutput power of its communication module. For this purpose, the servicevehicle preferrably is equipped with a control module for providing asetpoint for an output power of said configurable communication module.In further preferred embodiments, the setpoint for the output power ischosen in dependence of the current distance between service vehicle andtarget spacecraft and/or the relative orientation of the targetspacecraft with respect to the service vehicle. Accordingly, the controlmodule preferrably inputwise is connected to a first position sensor,said first postion sensor delivering a set of data characteristic forthe current position of said service vehicle, to a second positionsensor, said second position sensor delivering a set of datacharacteristic for the current position of said target spacecraft,and/or to an orientation sensor, said orientation sensor delivering aset of data characteristic for the current orientation of said targetspacecraft in relation to said service vehicle.

In a particularly advantageous embodiment, which may also be usedindependently from the communication concept as identified, the servicevehicle is designed for reliable and easy-to-use docking at the targetspacecraft. For this purpose, it preferably comprises a docking system,said docking system comprising a hollow first axle inside of which asecond axle is moveably disposed, said second axle carrying anactivateable arrow tip. For docking purposes, the activateable arrow tipmay be inserted into the exhaust system of the thrusters of the targetspacecraft via the axle system. Once inserted into the interior of theexhaust system, the arrow tip, preferably a double-arrow tip, may beactivated in order spread the arrow fingers apart. Retracting the arrowtip via the axle system will then cause the arrow tip to engage with theside walls of the engine exhaust, thus providing for reliable docking atthe target spacecraft.

With respect to the servicing system, the object identified above isachieved with a service vehicle as described above, further supplementedby a ground control unit for delivering operational commands to theservice vehicle. In order to consequently use the target spacecraft forrelaying communication signals from the service vehicle to groundcontrol in this servicing system, the ground control unit preferably isset up to receive data from the service vehicle by using the targetspacecraft as a relay station for signals emitted from the servicevehicle.

The servicing system may further be supplemented by an orbit-basedservice base for the service vehicle and/or by a propulsion moduleattachable to said service vehicle.

With respect to the method for in-space servicing of a selected targetspacecraft, the object identified above is achieved in that a servicevehicle as identified is used to perform selected in-space operations onthe target spacecraft, whereby operational signals from the servicevehicle are transmitted to a ground control unit by using the targetspacecraft as a relay station for the operational signals.

Alternatively or additionally the object identified above is achieved inthat a service vehicle as identified is used to perform selectedin-space operations on the target spacecraft, whereby a telemetrychannel between a ground control module and the target spacecraft isused to relay command signals to said service vehicle. In this setup,the telemetry channel which is auxiliarily used to echo telecommands canecho telecommands destined to service vehicle as well.

Among others, the main advantages of the present invention are thatparticularly inexpensive apparatus and methods for performingparticularly inexpensive science missions from space, requiringconsumables or/and robotic facilities, are provided. Furthermore,particularly inexpensive apparatus and methods for altering orbits ofpassive or active objects in space for whatever reason (anti-collision,operational) or maintaining its position against perturbing forces areprovided as well as inexpensive apparatus and methods for effectingreconfiguration, maintenance and/or assembly operations. Stillfurthermore, the invention pertains to reusable synergetic apparatus andmethods for performing inexpensively a variety of proximity operations,e.g., inspection of an operational or non-operational satellite, todetermine its status, (its weight, its temperature profile, theoperation or its subsystems), and/or to reusable synergetic apparatusand methods for inexpensively delivering or replenishing supplies toorbiting spacecraft or complexes such as the international spacestation.

Furthermore, to ground or elsewhere a high bandwidth telecommunicationlink originating from a simple inexpensive low powered servicing moduleis provided, optionally together with a simple method of controlling aspacecraft through part of the telemetry produced by another spacecraft,and/or an inexpensive apparatus and method for recovering telemetryinformation from a spacecraft whose telemetry means transmit at very lowpower. Still furthermore, the invention provides apparatus and methodfor recovering telemetry information from a spacecraft whose telemetrymeans transmit in very low power and encrypt it before retransmissionthrough on-board means or through means of the serviced spacecraft,and/or an inexpensive simplified mechanical grip for capturing asatellite from the interior of the combustion chamber of the satelliteand method of securing the grip, resulting to a pair of bodies(satellite & service module) of high stability.

With the receiver setup to communicate via the telemetry channel it isalso very easily possible to use telemetry data exchanged between thetarget spacecraft and service vehicle for real time diagnostics.

An in-space service vehicle, in order to provide even the minimum ofservices, namely inspection, it requires to be equipped with one or morecameras and means to establish an associated High BandwidthCommunication Link (HBCL) to the ground. Through this link it providesin real-time, the visual or infrared or other high bandwidth informationthat is captured, to teleoperators at the ground, to enableteleoperation. The said link requires very demanding resources (power,telecommunication means), especially if the service is to be offered atthe geostationary ring level.

The method in accordance with the invention includes usage oftelecommunication means of said satellit for the transmission of thesaid images to teleoperating controllers at the ground segment and noteffecting as usually has been proposed, the link directly to the GroundStations in an autonomous manner. The service vehicle proposed possessesmeans for transmitting at low power and at the frequency of anoperational up-link transponder of the target spacecraft 2 the videosignal properly modulated. The satellite shall retransmit as normallythe respective converted and amplified signal through the respectivedown-link transponder. Preferably, the up-link transponder of theoperational transponder chosen for the said link shall cease operationduring the service mission to allow unhindered image reception to theGround Segment.

An examplary embodiment of the present invention is explained in greaterdetail with reference to the drawings in which:

FIG. 1 shows a first version of a servicing system for providingin-space service operations to a selected target spacecraft,

FIG. 2 shows a second version of a servicing system for providingin-space service operations to a selected target spacecraft,

FIG. 3 shows a service vehicle of the servicing system according to FIG.1 or FIG. 2 docked to the target spacecraft,

FIG. 4 shows a schematic structure of the communication system of theservice vehicle according to FIG. 3,

FIG. 5 shows a utility base of the servicing system according to FIG. 1or FIG. 2,

FIGS. 6 a, b show a flexible storage module of the utility baseaccording to FIG. 5 in inflated (FIG. 6 a) and deflated (FIG. 6 b)condition, respectively,

FIG. 7 shows a schematic view of the internal layout of an equipment andstorage bay of the utility base according to FIG. 5,

FIGS. 8 a-c show a robotic manipulator for use in the interior of theequipment and storage bay according to FIG. 7 in side view (FIG. 8 a)and in top view (FIG. 8 b) and a cross section of a rail system for therobotic manipulator (FIG. 8 c),

FIG. 9 shows a docking and refueling rack of the utility base accordingto FIG. 5,

FIGS. 10 a, b show a side panel (FIG. 10 a) and a top panel (FIG. 10 b)of the docking and refueling rack according to FIG. 9,

FIG. 11 shows a catch system, particularly for use in the utility baseaccording to FIG. 5, and

FIG. 12 shows an action tip for the catch system according to FIG. 11.

In all figures, identical parts are provided with identical referencenumerals.

Following terms as used herein mean:

Spacecraft: is any type of manmade apparatus that is launched in spaceas a whole or produced through assembly in space.

Satellite: is a spacecraft that has entered a roughly determined orbitaround a celestial body (planet, natural satellite or sun). “Orbitalelements” are called the set of parameters that are describing thisorbit.

Delta velocity: is the velocity increment or decrease of a spacecraftwith respect to its vector of motion, by the application of a force thatis called trust and is provided through the thrusters of the spacecraft.

Total delta velocity potential: is the cumulative sum of the deltavelocity a spacecraft can generate throughout its operational life.

Geostationary object: is an object that has an eastwards circular orbitaround earth at a height of about 35,786.4 KM above the sea level.Characteristic of tremendous significance of this orbit is the fact thatthe object rotates with the same angular velocity as the earth and so itis visible as stable above the equator at certain Longitude, makingpossible the continuous communication with it through a single stablypointing antenna. The sub-satellite point is stable and is located at acertain longitude at the equator.

Station keeping maneuvers: are these maneuvers that are required to putor return a spacecraft to its desired point (or trajectory for missionswith no stable sub-satellite point eg Molniya) of operation.

Fail-Safe: a technical characteristic of an apparatus that is designedin such a way that when it fails for any reason it does not pose a riskapart from the loss of service it is designed to offer.

The servicing system 1 according to FIGS. 1 and 2 is designed to providein-space service operations to a selected target spacecraft 2, inparticular a target satellite, at both high reliability levels and lowfuel/cost levels. In this context, the sevicing system is designed toprovide the services both to so-called cooperative (or controllable)targets as is shown in FIG. 1, or to non cooperative (or noncontrollable) targets as is shown in FIG. 2.

In order to provide services in a broad variety of missions, typicallyin each mission type units of several, in particular three, species areused. These various species of spacecraft, in various numbers dependingupon mission, co-operate in a synergetic manner in order to provide aservice to the target spacecraft 2, either cooperative ornon-cooperative.

As a first element, the servicing system 1 comprises a module serving asa utility base 4, in the role of mothership for further elements. Thesecond element, a service vehicle 6, has the role of the actual serviceprovider to the target spacecraft 2 and may also be referred to as a“Utility Agent service vehicle 6”. A third element is an engine module8, potentially a subset of the service vehicle 6, suitable for permanentorbit maintenance service on a cooperative target. As fourth element, aspecialized vehicle 10 for missions with non-co-operative targets, orfor carrying and operating specialized repairing means or communicationrelay means, also referred to as “Escort Agent EA” may be provided.

By use of the servicing system 1, the existing fleet of spacecraft canbe adequately serviced and upgraded, and future spacecraft can beproduced at much lower cost, much lower mass and much shorter time,making use of the advanced repairing and upgrading capabilities of theservice fleet of the servicing system 1. Operational life of spacecraftis extended, capabilities are augmented, space debris problem ismitigated. In this context, the service vehicle 2 is designed to providea series of operations dissimilar in nature and complexity. In general,a single service vehicle that would embody all potential characteristicswould be too expensive to construct, as many studies have shown.Furthermore, its size and weight would increase the risk and operationalcost. Taking into account the potentially large variety of mission typesand that it would require to perform high and often changes in velocityany saving in weight budget would pay back many times.

Therefore, the service vehicle 6 is designed for particularweight-effectiveness and flexibility. This primary goal is achieved bythe fundamental design philosophy that it is specially designed to beteleoperated through a high bandwidth link via the target spacecraft 2itself. On that respect it remains autonomous from the utility base 4for long although small and it gains reusability potential by the meansof the utility base 4. Accordingly, in order to allow for low energyconsumption and the corresponding savings in weight (i.e. for energystorage devices such as batteries), the service vehicle 6 is designed tocommunicate with a ground control module 12 via a relay station. In theoperating mode as shown in FIG. 1, the target spacecraft 2 itself isused for relay purposes. As indicated by the arrows 14, 16, signalsemitted by the service vehicle 6 are transmitted to the targetspacecraft 2, thus according to close proximity demanding only limitedtransmission power. The service vehicle 6 emits the signals to thetarget spacecraft 2 in such a way that the target spacecraft 2 isoperated to forward the signals to the ground control module 12, forthis purpose providing the required (comparatively high) transmissionpower from ist onboard energy sources.

In case a non-cooperative target spacecraft 2 is to be serviced, as sisshown in FIG. 2, the service vehicle 6 may be accompanied by aspecialized vehicle 10 in this context providing the necessarytransmission power.

In order to facilitate using the target spacecraft 2 for the intendedrelaying purposes, the service vehicle 6 is equipped with acommunication module that can be configured such that it can communicatewith an arbitrary target spacecraft 2 in order to command it to forwardincoming signals to ground control module 12.

The service vehicle 6 is shown in more detail in a position docked tothe target spacecraft 2 in FIG. 3. Within an outer main body 20, aplurality of servicing facilities (not shown in detail, but selectedappropriately to provide the service required) is disposed. Attached tothe main body 20, there is a separable propulsion system 22 mainly basedon the use of conventional thrusters. In order to firmly attach itselfto the target spacecraft 2 after the final approach, the service vehicleis equipped with a docking system 24 designed to engage with the exhaustsystem 25 of the target spacecraft 2. In order to provide visualinformation for final approach, or to inspect the target spacecraft 2, anumber of cameras 26 is attached to the main body 20.

The service vehicle 6 is equipped with a built-in communication system28, which datawise is connected to an altitude and orbit control system30 of the service vehicle 6. The communication system 28 is designed to,at close enough distances, establish a communication channel with theso-called up-link communication channel of the target spacecraft 2. Forthis purpose, as indicated by the dashed line 32, the communicationsystem 28 establishes a communication channel with a receiver 34 of theup-link channel of the target spacecraft 2. Via this communicationchannel, the communication system 28 transmits commands or signals at acomparatively low output level that within the target spacecraft 2 arerelayed and forwarded to the emitter 36 of the so-called down-linkchannel of the target spacecraft 2. As indicated by the arrow 38, thesignals are then forwarded via the down-link channel to the groundcontrol module 12 at a comparatively high transmission power, the enegryfor which is derived from the on-board energy sources of the targetspacecraft 2.

For easier maneuvering relative to the target spacecraft 2, the servicevehicle 6 is equipped with a laser unit 39 set up to identify thedistance of the service vehicle 6 from the target spacecraft 2.

The docking system 24 of the service vehicle 6 mainly comprises a hollowaxle 40, an activation axle 42 inside the hollow axle driven by afail-safe mechanism 44 that allows extension, retracting or rotation ofthe hollow axle. At the free end of the activation axle 42, a doublearrow opening tip 46 (one arrow being smaller than the other) isprovided. The double arrow opening tip 46 is opening by means ofretracting the activation axle 42 and an even surface around theactivation axle 42 to permit even contact of the front surface 48 of theservice vehicle 6 with the nozzle ring 50 of the exhaust channel 52 ofthe target spacecraft 2, providing high stability when engaged.

The method of docking consists of the following phases: alignment ofaxle 40 to nozzle 50, entering the activation axle 42 into combustionchamber 54 of the target spacecraft 2, opening of the arrowheads,rotation if needed with stepwise retracting, final retracting of hollowaxle 40 and fail-safe engaging of the double arrow opening tip 46 withthe interior of the combustion chamber 54.

At approaching the target spacecraft 2, the arrow head sides shall bealigned parallel to the axle 40. The axle 40 is guided towards thecenter of the combustion chamber 54 through the nozzle 50 and when itpasses the neck of the chamber 54 the arrow head sides are opened wideto the maximum, through retracting the activation axle 42 in order tosecure it inside the combustion chamber 54. If the angular alignmentbetween service vehicle 6 and target spacecraft 2 is satisfactory thenthe securing and safeing phase is started, if not then the mechanism 44retracts the hollow axle 40 and rotates the activation axle 42 insuccessive steps until the desired angular alignment is achieved. Thenthe retreating mechanism 44 retreats slowly and firmly the hollow axle40 until the surface of the service vehicle 6 reaches and presses ontothe nozzle end-ring of the target spacecraft 2. The activation axle 42is fail-safe secured at this position and is released only by command orif a general failure occurs. In case of a power failure or mechanicalfailure or processing failure the activation axle 42 is left to itsnatural position by means of a spring that forces the arrowheads close.An independently powered timer controls the safeing mechanism. As longas the anomaly detection mechanism has detected no anomaly threateningthe target spacecraft 2, the activation axle 42 pushes open thearrowheads. In the case a threatening anomaly is detected the activationaxle 42 is left free and, forced by a spring, lets the arrowheads close.Any forward movement of the target spacecraft 2 lets the service vehicle6 to free float in space.

The structure of the communication system 28 of the service vehicle 6 isshown schematically in FIG. 4. As a key component, the communicationsystem comprises a communication module 60 which is designed such thatwith respect to its transmission characteristics it may be configured inorder to meet given receiver parameters of the selected targetspacecraft 2. Accordingly, by proper configuration of the communicationmodule 60, communication with any kind of target spacecraft 2 may beestablished and hence the service vehicle 6 can be teleoperated by usingthe target spacecraft 2 for relaying signals.

The communication module 60 comprises a multiplexer 62, connected to asignal modulator 64. Multiplexer 62 together with modulator 64 generatethe signals to be transmitted. For transmission purposes, thecommunication module 60 further comprises a transmitter 66 in connectionwith the modulator 64. For configurability, the transmitter 66 isequipped with a controller module 68, which if supplied with therequired data format may reconfigure the transmission characteristics ofthe transmitter 66 on a software basis. Furthermore, within thecommunication module 60, the transmitter 66 is exchangeable.Accordingly, configuration of the communication module 60 may also becarried out in a hardware manner by providing an alternative transmitter66. Since there are a plurality of satellite types or categories,preferable configuration is carried out on a hardware basis, i.e. byreplacing the transmitter 66, if reconfiguration bewteen differenttarget spacecraft categories is desired, whereas reconfiguration is doneon a software basis, i.e. by reprogramming the controller module 68, ifreconfiguration between different individual target spacecraft of thesame category is desired.

Inputwise, the multiplexer 62 is connected to an encoder 70, which inturn receives its input data from a camera 72 and/or a proximity sensor74. Furthermore, the multiplexer 62 inputwise is also connected to atelemetry system as indicated by the arrow 76.

With respect to its output power, the transmitter 66 is adjustable inorder to make sure that the power emitted will not endanger or destroythe target spacecraft 2 due to close proximity. Accordingly, thetransmitter 66 is equipped with a control module 78 designed to providean appropriate setpoint for the output power. The control modulepreferably generates the setpoint for the output power based upon asignal strength received from the target spacecraft 2, which ischaracteristic for the relative distance of the service vehicle 6 fromthe target spacecraft 2. Accordingly, inputwise the control module 78 isconnected to a communication receiver 80 of the communication system 28.The receiver 80, which inputwise receives signals from the targetspacecraft 2 as indicated by the arrow 82, outputwise is connected togeneral data handling of the service vehicle 6 via a demodulator 84.Further components, such as a docking subsystem 86, the proximity sensor74 directly via a branch line 88, retroreflectors 90 mainly used forother spacecraft to dock on, or an optional refueling module 92 are alsoconnected to a telecomand bus or general data handling of the servicevehicle 6.

The receiver 80 is adjustable in its working frequency in order tocommunicate with a telemetry channel of the target spacecraft 2.Accordingly, by ground setup of the communication parameters, esp.frequency, of the receiver 80, control or command signals may be sent tothe receiver 80 by using the so-called telemetry echo channel of thetarget spacecraft 2. In this configuration, command or control signalsfor the service vehicle 6 may be included into the data stream sent tothe target spacecraft 2 in the conventional telementry channel.Preferably in this case the command or control signals are provided withan associated identification element or tag. In the target spacecraftthe data elements identified in this way can be re-emitted in thetelemetry echo channel to be picked up by the receiver 80. In accordancewith their identification, these signals can then be forwarded forfurther processing in the service vehicle 6. This specific setup inprinciple makes if possible to design the service vehicle 6, which inparticular may be a utility agent or an engine module, without aseperate navigation system since the entire telemetry information fromthe spacecraft 2 may be forwarded to the service vehicle 6.

Beyond, the functional composition of the bus system of the servicevehicle 6 comprises the following subsystems: a structure subsystem, thedata handling subsystem (DHSS), an electric power subsystem (EPS), athermal control subsystem (Ttarget spacecraft 2), an attitude orbit &control subsystem (AOtarget spacecraft 2), a telemetry tracking &control subsystem (TT&C), and a propulsion subsystem (PSS),characterized by no redundancy in any of the subsystems budgets.

Albeit the fact that these subsystems are present in the majority ofspacecrafts the bus of the service vehicle 6 is characterized by lowcapability budgets of the respective subsystems, in account of itsmission and the lack of redundancy. The lack of redundancy is justifiedby the capability, in case of failure of a given fleet unit, ofrecovering it through another service vehicle 6 or specialized vehicle10 and subsequently repairing it at the utility base 4.

In particular, the EPS consists of small solar cell array panels (SAP)capable to produce part of the energy required during missions. Start ofmission charging is performed at the utility base 4 before the missionstarts. Likewise, the batteries of the service vehicle 6 are undersized,as at proximity to the utility base 4 the telemetry is relayed throughthe utility base 4, at cruise if needed directly to earth and then atapproach of the target spacecraft 2 through the target spacecraft 2. Atproximity to the target spacecraft 2, the target spacecraft 2 is used asrelay for both the TT&C and the cameras output. The EPS does not caterfor any high-bandwidth link to support teleoperation or robotic facilityor both as it is usually being proposed. Considering that the EPS of atypical spacecraft is 30% of its mass budget this saving is of highimportance.

The TT&C transmitter is of low bit-rate and characterized by the use ofAdaptive Power Control APC. The TT&C transponders can be switched offwhen in proximity to the target spacecraft 2. In this case the telemetryTM and telecommand TC are transferred through the payload.

The service vehicle 6 to perform docking and operations establishes oneforward link with the teleoperators, preferably at ground control module12, and a return link both through the target spacecraft 2.

The forward link is established as follows: The encoder 70 of theservice vehicle 6 payload receives two inputs, one for the signal of thecamera 72 and one for the proximity sensor 74 and generates two encodedsignals for the camera signal and the proximity sensor respectively. Themultiplexer 62 receives these two signals plus the encoded TM signalfrom the DHSS of the bus and multiplexes the three, producing acomposite signal. The modulator 64 receives the composite signal,produces a modulated signal and feeds the transmitter 66 which amplifiesand transmits the signal that is fed to the up-link receiver of achannel of the target spacecraft 2. The target spacecraft 2 receives thesignal and transmits to the ground. The transmitted signal arrivesthrough the ground control module 12 at a Mission Control Centre (MCC)for analysis and informed action.

The teleoperators in the MCC generate telecommands for the servicevehicle 6, which are embedded within the telecommands for the targetspacecraft 2. These telecommands for the service vehicle 6 are flaggedwith the request only to echo them and not to be executed by the targetspacecraft 2. Following the reception of the telecommands the targetspacecraft 2 echoes them from the telemetry channel. This signal iseasily intercepted by telemetry receiver of the service vehicle 6.

The telecommand reception is established as follows: The telemetrylisten-in receiver receives the totality of the telemetry of the targetspacecraft 2 and produces a signal that forwards for demodulation at thedemodulator 84. After demodulation the resulting signal is forwarded tothe DHSS of the bus and in particular at the application software wherethe analysis of telemetry is performed for extracting this informationthat consists commands to the service vehicle 6.

The main types of operation of the service vehicle 6 in relation with amission are cruising from the utility base 4 which is serving as astarting platform for each mission, approaching the target spacecraft 2(rendezvous and teleoperation), return from the target spacecraft 2 tothe utility base 4, and resting at the utility base 4 until the nextmission for the respective service vehicle 6 is started.

When cruising from the utility base 4 to the target spacecraft 2(“Cruise mode”), the service vehicle 6 travels from the utility base 4to the target spacecraft 2 alone and autonomously making use of the startracker. The power output of the TT&C of the bus is adjusted so thattelemetry link can be established by the bus TT&C through either theutility base 4 or the target spacecraft 2. If neither is possible due tolarge distances, the service vehicle 6 may be escorted in the neededpart of its cruise by a specialized vehicle 10, may be used to relaytelemetry and telecommands from a ground control module 12 to theservice vehicle 6 and vice-versa, thus rendering the service vehicle 6operable in any state of the cruise in spite of its limited on-boardtransmission and fuel capacities.

For rendezvous and teleoperation, during the coast phase from theutility base 4 to the proximity of the target spacecraft 2 the starimages from the cameras 26 are used for autonomous navigation. Duringthe approach and rendezvous phases the service vehicle 6 is controlledby means of open loop successive command cycles until docking issecured.

At each command cycle the real-time output of the cameras 26 is encoded,multiplexed, and modulated together with telemetry information of theservice vehicle 6 (and optionally with the output of the proximitysensor 74). The resulting signal is transmitted by the low powertransmitter 66 to an up-link channel of the target spacecraft 2 throughits up-link antenna. The target spacecraft 2 retransmits through therespective down link channel said signal to the ground control module 12which may be part of a ground station (GS) and mission control center(MCC). The receiver of the ground controil unit 12 receives thecomposite signal, demodulates and de-multiplexes and then decodes theimage, telemetry and proximity sensor signals and forwards them to theMCC. The telemetry information and proximity sensor information isrecorded at the MCC, analyzed and several derivative parameters aregenerated to optimize motion commands of the teleoperation apparatus.Said optimization compensates for fuel mass changes, sloshing activity,thruster efficiency, fuel temperature, combustion chamber temperatureand other biasing factors difficult to be handled by an operator in realtime. The real-time image together with the summary proximityinformation and other rendezvous related information (relative angles,time windows of critical steps, fuel reserves etc) is displayed ontovirtual-reality head-on display systems of a plurality of teleoperators.

Said teleoperators have control over actuators generating appropriatecommands which pass through the above said optimization. Said optimizedtelecommands are packed in special telecommands of the target spacecraft2 and are forwarded from the MCC to the transmitting part of the groundcontrol module 12, encoded, modulated and transmitted as part of thetelecommand stream to the target spacecraft 2 with appropriateidentification. The telecommands that are addressed to the servicevehicle 6 are echoed by the down link (telemetry) of the TT&C of thetarget spacecraft 2 and listened-in by the TT&C receiver of the servicevehicle 6. The listened-in telemetry signal is demodulated and decodedand a telecommand selector parses the telemetry and selects telecommandsaddressed to the service vehicle 6. The said telecommands are executedand telemetry is generated that in turn is encoded, multiplexed with theoutputs of the cameras 26 and proximity sensor 74, modulated and thentransmitted to the selected up-link channel of the target spacecraft 2.

This command cycle is repeated until the docking system 24 is securelyfastened inside the combustion chamber 54 of the target spacecraft 2.

Upon mission completion or fuel shortage, the service vehicle 6 returnsto the utility base 4 for resting or refueling, respectively.

In proximity to the utility base 4, maneuvering of the service vehicle 6is assisted by the surveillance means of the utility base 4. The servicevehicle 6 assisted by the utility base 4 sensors and retroreflectorsperforms preferably an automatic docking at the utility base 4. However,teleoperated docking may also be performed.

In the “resting mode”, under service-call wait-status, the servicevehicle 6 rests, preferably at the utility base 4, preferably in astorage mode that consumes very limited resources. It is envisaged that,at full deployment, there will be provided a multitude of servicevehicles 6 at a single utility base 4 with some variations in size andinterfaces to correspond to specific types or categories of targetspacecraft 2, or to better a match a selected type or level of serviceto be provided to the target spacecraft 2.

In case that the target spacecraft 2 requires specific services fromsubsystems of the utility base 4 (robotic facility), the service vehicle6 may be operated to fetch the target spacecraft 2 to the utility base 4for servicing and places back to the desired post after service(“porting mode”).

The service vehicle 6 depending of the mission duration may be equippedwith additional fuel reserves and a fuel delivery subsystem. In anothervariation, the service vehicle 6 may be designed to perform a variety ofmissions with add-on accessories. For example, a service vehicle 6equipped with drilling means and endoscope may be used in tandem with aspecialized vehicle 10 for performing indepth investigations of failurecauses or other rescue missions.

The engine module 8 of the service vehicle 6 primarily is used for orbitmaintenance of a target spacecraft 2 and for potentially reserving fuelof a target spacecraft 2. The engine module 8 comprises a subset ofelements of the service vehicle 6. In particular, the bus of the enginemodule 8 may be part of the attitude and orbit control subsystem if themission is propulsion only. Its payload consists of a fail-safedocking-securing mechanism identical with the one of the service vehicle6 and a TT&C that interfaces with the TT&C of the target spacecraft 2 ina way similar to the concept of the service vehicle 6. This TT&Ccomprises a telemetry listen-in receiver-demodulator-decoder-commandselector and an encoder-modulator-transmitter that transmits to theup-link of the TT&C channel or other channel, as preferably of thetarget spacecraft 2.

The engine module 8 possesses electrical and data interfaces for matingwith a porting service vehicle 6, and optionally a fuel reception inlet.It disposes at all sides retroreflectors that facilitate automaticdocking of a visiting or refueling service vehicle 6. The engine module8 may be used to be forwarded and attached to a target spacecraft 2 bymeans of a service vehicle 6. When mission fuel depletes it receivesadditional fuel by a refueling service vehicle 6. Return to the utilitybase 4 may then require a porting service vehicle 6. In case of criticalfailure the fail-safe mechanism is automatically released.

The level of redundancy of the engine module 8 is customizable accordingto mission requests. An engine module 8 for a target spacecraft 2 withno fuel reserves preferably has full redundancy. An engine module 8 fora target spacecraft 2 with fuel sufficient for a few months operationmay be designed with no redundancy.

At full-scale deployment of the servicing system 1, a plurality ofutility bases 4 may be held available. The most preferable position tostart with is the geostationary ring, less preferable the low earthsunsynchronous polar orbit. Any other possible orbital plane is objectfor positioning a utility base 4 but markets other than that of thegeostationary ring and the sunsynchronous polar orbits need still to bematured.

The utility base 4, which is shown in FIG. 5 in more detail, representsthe mother ship for service vehicles 6 or other vehicles 10 of theservicing systen 1. As main components, the utility base 4 comprises amain body 100, which primarily houses control systems and the like andcontains the bus system of the utility base 4, an equipment storage bay102, a docking/refueling rack 104, and a flexible storage module 106.The interfaces between these segments dispose power, data “TMTC” andplurality of video signal connectors.

Attached to the main body 100, primary solar panels 108 are provided forenergy supply. For redundancy purposes, secondary solar panels 110 areattached to the equipment/storage bay 102. The equipment/storage bay 102further carries a support grid 112 for securing and storing items ifneeded. In order to potentially move items around, a robotic arm 114preferably extending beyond the support grid 112 is mounted onto themain body 100. For establishing communication channels, a number ofreflectors 116 of antenna are attached to the equipment/storage bay 102.The primary and redundant large aperture parabolic antennas are mountedonto the down-out side of the equipment/storage bay 102.

In order to allow for docking of a multitude of service vehicles 6 orspecialized vehicles 10, especially for resting purposes without theneed for supplying the respective vehicle further, the utility base 4 isequipped with a number of docking stations 118. Although in FIG. 5 onlyone docking station 118 is explicitly identified, further dockingstations (preferably at least four in total) are provided, preferably atleast one in every main direction of the utility base 4.

In general, the utility base 4 is characterized by a “hot redundant”architecture protecting against two points of failure of all its vitalfunctions (links to the ground, robotic functions, docking spaces) andmechanisms (e.g. electric power subsystem, attitude control subsystem),providing survivability of itself and of the carrying fleet againstdouble failures.

The utility base 4 comprises means of active and passive surveillance ofthe surrounding space (ranging lasers, radar systems) and has activemeans (potentially reying on docked or otherwise available servicevehicles 6) for avoiding collisions with other elements in open space(ablating laser). Given the replenishment capability of its resourcesthrough often replenishment missions and the high redundancy of is vitalfunctions, the utility base 4 that is placed at the geostationary ringmay in essence be the first space platform with indeterminable lifespan.

It is used to perform surveillance, protection, positioning, hosting,storing, reconfiguring, repairing, converting, assembling, and sciencemissions.

Assuming the position of the utility base 4 at the Geostationary ring atmid day, a coordinate system passing from the geometric centre of itscentral segment is defined as follows. X axis has west to eastdirection, Y axis has Earth to Sun direction and the Z axis has South toNorth direction. For the X axis also the left-right notion is used whereX increases to the left, for the Y axis the near-far notions are usedwhere Y increases towards far, and for the Z axis Up an-Down notions areused where Z increases towards up direction. When relative reference ofa segment of the utility base 4 other that the central one is made, inrelation to the centre of the utility base 4, the terms IN-side andOUT-side are also used. In-side denotes the side close to the centre andout-side meaning the side of the segment at question which is oppositeto the In-side at a direction departing from the centre.

The bus system of the utility base 4 mainly consists of a doubleredundant TT&C subsystem, a redundant attitude and orbit controlsubsystem (AOCS), a redundant electric power subsystem (EPS), aredundant data handling subsystem, and a redundant thermal controlsubsystem (TCS). All subsystems are characterized by hot redundancy. Theutility base 4 receives power primarily from the solar panels 108(preferably three or more) mounted onto booms connected to an axialtruss through mechanisms having three degrees of freedom. The truss ischaracterized by passing from the geometric and momentum center of themain body 100 through the same axis as the robotic arm 114. Theactuators of the solar panel mounting mechanisms of the primary andredundant solar panels 108, 110 are part of the AOCS.

The robotic arm 114 is designed to have five degrees of freedom (DOF)for the actual arm 120 and three degrees of freedom for its wristelement 122. The robotic arm 114 is dimensioned such that it can reachall upper, side and under areas of the utility base 4 that may needservicing.

The communication system or payload of the main body 100 also possessesa redundant near range mission communication system, preferably forten-channel RF video reception equipment, a video switch system, and aredundant communication payload, for transmission to the ground of fouruncompressed and twelve compressed digital video signals, generated bythe various surveillance and teleoperation cameras. The redundancy ofthe mission communication system to the ground may provided by aspecialized vehicle 10 docking at the far end of the equipment/storagebay 102.

The utility base 4 does not necessarily possess its own propulsionsystem, but fleet units (service vehicles 6/specialized vehicles 10) maybe attached to the four sides and commanded appropriately when neededfor orbit maintenance. Attitude stability of the utility base 4 isachieved, in short time, by use of the steering mechanisms of the solarpanels 108,110. The utility base 4 is axi-symmetrically momentumstabilized.

The flexible storage module 106 mainly consists of a flexible,inflatable, lightweight balloon-like surface sheet, the size and shapeof which may be modified by retreating means 124. In the embodimentshown, the retreating means 124 mainly are provided by contractabletapes which when contracted will diminish the volume of the interior ofthe module 106 while increasing its volume when allowed to expand.Examples for the module 106 in expanded and in contracted status areshown in FIGS. 6 a and 6 b, respectively. Accordingly, the flexiblestorage module 106 resembles a sack-shaped flexible storage bay whichpossesses a plurality of ring shaped, tape-measure type tape-fastener,externally secured to the sack by means of externally to the sacksecured small elliptic fasteners. Said ring tape is driven by areel-unreel mechanism with dual reels having independent motors. Byreeling-in the tape the sack closes securing the free flying objectsthat are placed in this sack and by unreeling the tape the sacks opensto let the robotic arm 114 or other means collect the objects. Anothertape fastened perpendicular to a securing ring on the external surfaceof the sack elongates or shortens the sack respectively, increasing ordecreasing its volume.

The equipment/storage bay 102, the interior of which is schematicallyshown in FIG. 7, and which also may be referred to as a closed equipmentstorage bay (CESB), is mainly used for housing equipment and materialsensitive to exposure to radiation, or temperature variations, orsun-rays, or small meteorites. It is where assembly, disassembly andtesting takes place for small mechanical, electromechanical orelectronic subsystems. The treatment of the material to be handled mayor may not include packaging and un-packaging in protective boxes.

The west side of the equipment/storage bay 102 disposes a pressurizationcontrolled pro-thalamus 130 with five outer doors 132 and a singleinternal door 134. The west door and inner door 134 are disposed oneopposite to the other in a way to allow long objects equal to the longaxis of the chamber to enter the bay in unpressurized conditions.

The equipment/storage bay 102 possesses conditioning means for effectingand controlling pressure, temperature and cleanliness by Nitrogen gas orother inert and nonvolatile gas. It possesses permanent cameraviewpoints, equipment bay for manipulation of miniature mechanisms andelectronic circuit boards and components.

The up-side and down-side in the thalamus 130 for further descriptionare defined with respect to the position of the horizontal axis, upbeing the position where lighting sources and gas in-jets are mounted,down being the position where gas outlets are mounted. The gas jets arespread all along ceiling and gas outlets all along floor surface. Theflow of gas from up to down creates a small pressure potential to thefree flying objects in a way similar to gravity.

Manipulation of movable equipment within the equipment/storage bay 102is performed by means of a number of three-arm small-sized robots 140slidably and rotatably mounted on two horizontally secured axis 142. Thelong axis of the equipment/storage bay 102 defines the horizontaldimension. A third axis 144 with an H profile, the profile of which isshown in FIG. 8 c, is disposed in between the above two mentioned axisand disposes two conductive surfaces 146 on its interior. Saidconductive surfaces 146 are used by a the robots 140 to slide alongwhile at the same time supplying them with electric power.

As shown in FIGS. 8 a, 8 b in greater detail, each robot 140 consists ofa pair of two cooperative human-like manipulation arms 148, each havingsix degrees of freedom, and a third arm 150 of three degrees of freedomthat is used for stability with a two finger gripper 152 designed to beengaged with the axis 144. Alternatively, for holding objects athree-finger gripper may be provided. The arms 148 of the robots 140have ten finger grippers each. The robots 140 can be positioned in aface-to-face configuration for cooperative work. The human-like arms 148of the robots 140 can be engaged to closed-chain kinematic configurationfor manipulation of objects. This means the one arm 148 follows intandem the movements of the other (driving) arm 148.

The robots 140 may be assisted by a plurality (minimum 2) of miniature(scale 1:3 of robots 140 or better) three arm robots 149 similar butwithout the sliding-rotation part of the robots 140. Mobility isprovided by a sliding mechanism perpendicular to the first element ofthe stability arm. With small jumping movements, using the two or threearms, the robots 149 can always reach a horizontal axis, attach thesliding mechanism of the stability arm and slide along. These robots 149either work from an axis or reach working place by a jump from theslide-on axis or are placed to workplaces by the robots 140. The robots149 are secured, when in workplace, by means of using their stabilityarm (with 3 degrees of freedom). Alternatively, they can be held by theholding arm of a robot 140 for common manipulation of an object inparallel, assuming the object is secured in place by other means. Therobots 149 when in workplace are connected to power/data/video-outputinterface and when in free float they use onboard power (batteries).Nevertheless, the floating time is limited and the respective batterysize accordingly. The robots 149 dispose accelerometers and gyroscopicmeans for attitude control when in free floating conditions.

The equipment/storage bay 102 disposes its further elements mainlyaround at mid level a bench surface, filled with holes for letting airpass through and create a small virtual gravity effect, and a stiff edgefor giving stability to the robots 140 when they grip on it. Disposesalso a plurality of grips for securing objects in place formanipulation. It further disposes a table 154 for common, face to facemanipulation with similar stiff edge, and a plurality of storage racks156 for storing/affixing tools, accessories, and spares. The stiff edgeand other places at the racks 156 possess connectors for providing therobots 149 with power/data/video interface. The distance of the storageracks 156 allows the robots 149 to use the stability arm to attachitself to a rack 156 while the other might be engaged to fetch/storeactivities. For moving from one rack 156 to another the robot 149 needsto stabilize itself by using the human like arms, gripping a horizontalshelf or a number of vertical bars, or a combination of a bar and ashelf, before disengaging the stability arm to move to another shelf.

The common table 154 is surrounded by tool & parts affix area mainly formechanical works and a tool & parts affix area mainly for electrical &electronic works.

The docking/refueling rack 104, which in further detail is shown in FIG.9, is designed to be semiautonomous and usable for all types of fleetvehicles 10, service vehicles 6, and the like. It is provided withstandardized utility outlets 160 for power, data, video, fuel, oxidizerand pressurization gas. At least two of the docking positions defined bythe outlets and their respective fixation means possess also relievein-lets for emptying the supplies of a fleet unit. Said inlets for fuel,pressurization gas, and oxidizer are disposed symmetrically to theoutlets, in respect to the docking unit centre. The docking/refuelingrack 104 has a plurality of pairs of docking interfaces for the fuel,oxidizer and gas tanks 162 (min two for each species), disposed at theupper and if needed also lower sides of the same. Each fleet unitdocking position has a pair of active securing mechanisms disposedsymmetrically to the centre of same. The tank docking positions haveeach a three-point active securing mechanisms. The schematics of theselocking mechanisms are shown in FIGS. 10 a, 10 b, which display the sidesurface 166 (FIG. 10 a) and the upper surface 168 (FIG. 10 b) of therack 104 with the other parts (esp. tanks 162) removed.

All fleet unit docking positions dispose retroreflectors for aidingapproach and docking. The centre of each fleet unit docking position ishollow to allow the grapple arrow pass the rack surface and secure theposition by opening the arrowheads and retracting.

Distributed pairs of docking positions without fuelling outlets but withdata and power outlets are disposed at all four sides of the utilitybase 4.

The docking/refueling rack 104 is semiautonomous in the sense that itpossesses a limited power supply storage system, a thermal controlsubsystem and a data handling subsystem that is designed for supportingdocking, fuelling operations and conditioning independently of the mainbody 100. The docking/refueling rack 104 can provide, through a datainterface, to the main body 100 of the utility base 4 all locallyavailable data.

A further position on the docking/refueling rack 104 is reserved for aspecialized vehicle 10 which can activate its cameras when needed, tosurvey the docking/refueling rack 104 and the rest of the utility base4. The video signal of the cameras can become available to the videoswitch of the main body either through a video interface or via RFtransmission to the RF reception payload of the main body 100. Thedocking/refueling rack 104 also possesses a redundant pressure-upequipment for helium gas which is operated only when connected throughthe interface to the main body 100. This capability of autonomousoperation allows for the disconnection of the docking/refueling rack 104from the utility base 4 when deemed there is increased risk associatedto performing hazardous operations such as refueling. Thedocking/refueling rack 104 in this case is removed by means of operatingone or more fleet units and is returned back when hazardous operationshave been completed.

The mechanical interface 170 that connects the dockin/refueling rack 104to the main body 100 disposes also connectors for the realization ofconnecting the various interfaces of the docking/refueling rack 104 tothe main body 100 (power, data, video).

Docking of other vehicles/objects is performed through customization ofextension constructs. After a target spacecraft 2 or another floatingobject towed by fleet units is delivered to the robotic arm 114 forstabilization, stabilization grids are erected as required for securingthe object in place and release the robotic arm 114 for otheractivities. These grids are constructed by means of a plurality of boomsthat are secured along the top of the equipment/storage bay 102, bymeans of fasteners.

Furthermore, the utility base 4 may be equipped with an open storage bay(OSB). Said bay is used to store equipment, tools, materials, productsand spares that do not require protection or conditioning, packaged orun-packaged. It may consist of two symmetric racks, east and west, whichare attached to the near side of the main body 100, through respectivemechanical, electrical, data, and video interfaces. Both racks (forredundancy purposes) comprise interfaces for operating (command, data)an externally mounted detachable parabolic antenna each, forcommunication with the fleet. In the case the stabilization grid isdeployed the redundant antenna is mounted onto the most western boom.They also both, for redundancy purposes, dispose interface for powercontrol and video for driving a catch system as will be explained below.The two racks are stabilized by means of a bridge 172 connecting theirnear sides. Said bridge 172 disposes in its middle a docking station 118for a fleet unit, preferably a service vehicle 6 or a specializedvehicle 10, which possesses cameras, and a shaft for mounting the catchsystem. The cameras of the service vehicle 6 or the specialized vehicle10 can assist fetching storing operations of the robotic arm 114 and ofthe catch system. The down inner corners of the storage racks, the downnear corner of the main body 100 and the down part of the rackconnecting bridge 172 dispose fastening points, respectively.

The catch system 180, which may be placed in different positions at theutility base 4, is shown in FIG. 11. Designed as a tape based capturetool (TCT), it mainly consists of a double reel-unreel mechanism 182mounted on a 3 degree of freedom mechanism (184), two conductive tapes(186) that extend in parallel, and an end piece 188. The end piece 188,which is shown in more detail in FIG. 12, is equipped with a camera, anumber of light sources, a 3 degree of freedom gripping wrist 190serving as a capturing mechanism. The catch system may be mounted onto adocking base sliding on a shaft attached centrally on the inside of therack connecting bridge 172, in a way that the cameras of the fleet unit(service vehicle 6 or specialized vehicle 10) docked on the bridge 172can supervise the activities of the same.

The catch system 180 is detachable from the docking base. Similardocking positions are available at the pressurized compartment of theequipment/storage bay 102 and on the far side of the open equipment bay.The robotic arm 114 can also capture and operate the catch system 180.The end piece 188 further possesses tension sensors for each tape,gyroscopic accelerometer 192, zero to four momentum wheels 194 forattitude control, RF means for transmission of the camera video signal,and a power conversion box 196. The power (alternating current) arrivesto the end piece 188 by means of the two conductive tapes 186. It isconverted to appropriate voltage ratings and distributed where needed.Control signals arrive to the end piece 188 by means of modulating thealternating current transported through the tapes 186. Video link istransmitted form the end piece 188 by means of an RF transmission. TheRF signals are received by a central RF reception base.

Small and medium volume objects for storage may be placed into boxes andboxes are secured in a set of adjacent shelves of parallelogram shape ofvarious sizes assembled out of aluminum or carbon fiber elements orother strong lightweight material. Said shelves may comprise a pluralityof temporal adhesive tags at their bottom side that secure boxes when inplace, even if a plurality of small boxes is stored into a large shelf.The fetching and storing of objects is performed by means of the roboticarm 114, the catch system 180, or other.

The upper side door 132 of the pro-thalamus 130 (FIG. 5) is reachable bythe robotic arm 114 and two appropriately positioned catch systems 180.All 5 outer doors 132 have mating interfaces for extension modules. Thepro-thalamus 130 houses a round rotating plate equipped with a catchsystem 180 in the one side of the table, which table can be raised, whenan outer door 132 of the pro-thalamus 130 is open, above the uppersurface of the equipment/storage bay 102. This way, an object that hasbeen placed on the pro-thalamus table with the help of the catch system180 can become available to the outside and vice-versa. The catch system180 can also make available objects to the interior of the main thalamusof the equipment/storage bay 102 when inner door of pro-thalamus 130 isopen.

In general, the fleet units of the servicing system 1, in particular theservice vehicles 6, do not have redundancy or means for significantlyreconfiguring themselves, as regards their hardware. Reconfiguration,repairing, assembling, upgrading is performed at the utility base 4using special purpose facilities. In addition, the upgrading subsystemis used for conversion of captured foreign objects (CFO). Said CFOs thatare of main interest for conversion are non-functional satellites, tanksfrom spent upper stages, and the like.

The upgrading subsystem comprises at least: an open equipment bay (OEB)and a protected, or closed equipment-storage bay 102 (CESB). Said OEB ismounted at the far side of the main body 100, through a mechanicalelectrical and data interface and the CESB is housed in a nitrogen gaspressurized chamber mounted at the west side of the main body 100.

Said Open Equipment bay “OEB” is used for mechanical or electrical workson the fleet, target spacecraft 2 s, or CFO. Conversion operations, bebetween else processes for effecting access windows on tanks, pipeconnecting/disconnecting, rack mounting, equipment and cabling networkinstallation.

Said OEB possesses a plurality of (minimum two) of human size dualrobotic arms (primary and redundant) for tool/manipulation with tenfinger grips, and arm articulation similar to the human (six degrees offreedom). Said dual robotic arms are movable on top of the main body 100and OEB by means of a mobile base that slides onto a T shaped rail pathmounted on their surfaces. The rail path starts at the near edge of theupper surface of the main body 100, crosses the upper surface of themain body 100 with direction towards the OEB. It passes at a sufficientdistance from the centre of the main body 100 where the robotic arm 114is mounted. Said rail path then crosses the OEB in a parabolic shape andthen passes on top of the CESB having a mounting point on it andcontinuing in a hemicyclic shape arriving to the upper side of thestorage rack.

Each robotic mobile base is driven by four powered wheels mounted onaxis parallel to the rail shaft and pressing against said T rail shaft.Six ball bearings for sliding along the rail head are provided as wellas four short ones mounted just below and two wide ones above the T railhead, mounted in parallel to the said horizontal T rail head.

OEB also possesses a plurality of tools and benches for performing thesaid services similar to what is found in the Ground Segment Supportequipment and particularly those that can be exposed to the open spaceenvironment with limited shielding.

The utility base 4 has a stock of accessories for repairing & upgradingthe fleet and own subsystems.

These accessories between else include replacement modules for the hotredundant elements of the utility base 4, (EPS, AOCS, MCP, RF, TT&C)telecommunication modules for UHF and S band and data channeltelecommunication modules for C, Ku and Ka band of various output powerratings. They further include attitude control sensors (sun, earth, starbased), cameras of various aperture ratings, filters, lenses, endoscopesand telescopic probes, towing tethers tether/wire deployment/retractingadd-on module as well as sets of retroreflectors, laser diodes, motors,ball bearings, lubricants and lubricating means. Adhesive materials,insulated wires, solar cell spares and fly wheel spares, valves andpipes, thrusters and any other accessory that may be foreseen, needassessment based on a statistical estimation of failure risks of thetarget spacecraft 2 components and subsystems.

Said repairing and upgrading tools comprising, between else, of hardwaretools set, (lathe, aluminum soldering, etc), electrical tools set (wireconnectors, soldering etc), electronic tools set (polymeters,palmographs etc.)

A plurality of tether equipped truss assists in the disassembly processby displacing disassembled elements away of the OEB core. Each time adisassembled element is attached to the tether the tether is promotedproportionally to the size of the attached element. To fetch a storedelement from the tethered truss the tether is advanced or retractedaccordingly.

The utility base 4 also is equipped with active and passive surveillancemeans.

These means are used for accurate positioning of objects in thesurrounding space and for protection from space debris as well as forassisting cruise or automatic docking of the fleet units. The proximityradar provides a coarse but wide image of the surrounding space objectsand the ranging laser a precise determination of distance and positionof objects in the surrounding space. The ablating laser destroys smallobjects or alters the trajectory of larger objects to avoid collisionwith target spacecraft 2 or utility base 4 or fleet units. It alsodestroys or steers the particles that escape from the manufacturingprocesses to a desired collection point.

The utility base 4 requires numerous video and Telemetry links to beestablished for full operation. A gradual process is envisaged toprovide the required bandwidth with use also of a resurrected satellite.

The specialized vehicle 10 may be designed to perform several functionsof a so-called escort agent (EA). It basically has the same functionalelements in its bus as a typical service vehicle 6 but reinforced interms of EPS budget and size. It is mainly used for missions with FCOand non-cooperative target spacecraft 2, or with target spacecraft 2where compatibility with its payload has not been achieved.

Its payload consists of two steerable high gain antennas, forestablishing receiving link and retransmitting link to differentdirections, and cameras. It is designed to assist the docking and otherservices of a service vehicle 6 by establishing the requiredsurveillance and teleoperation video links with a ground control unit 12directly or through the utility base 4, or through a third spacecraft.It receives through RF video and TTC signals from a service vehicle 6 ordirectly from its own cameras and retransmits after amplification.

A type of escort agents with refueling capability is defined for refugeerescue missions or other high energy orbit missions.

The primary operational concept for the servicing system 1 is to reusethe service vehicles 6 and other elements of the system in manymissions, servicing satellites that are far away in terms of deltavelocity potential required to reach them and carry them or maintaintheir orbit or optimize their trajectory, in particular by using thetarget spacecraft 2 for relaying signals to ground control.

Nowadays, most of the satellites are operating in the C, Ku and Kabands. Constructing communication means of very low power in a wide partof these bands to allow compatibility with a large population ofsatellites is not a problem. In addition to that, the utility base 4comprises means for performing extensive reconfiguration andcommunication module exchanges so that the service vehicle 6 can becomecompatible with almost the totality of the current satellite population.Since in short distances of a few meters to hundred meters away from thetarget spacecraft 2, the service vehicle 6 will have to operate the saidlink, directionality of the antennas is not that important and thatthere are backwards electromagnetic wave lobes that can be exploited forthis cause.

The advantage of the method is the provision of the needed bandwidthwith extremely low powered means. In the case where the powerfulcommunication means of the target spacecraft 2 are used as relay means,the means required in the ground for reception of the service vehicle 6is as simple as a simple TV receiver in the case of TV satellites.

Alternatively as it is foreseen in the case where the target spacecraft2 can not provide the required transmission means another specializedvehicle 10 will perform the task of establishing the link to the grounddirectly or through a relay, acting as relay satellite in the veryvicinity. In this case it might also observe the service vehicle by itsown means and provide alternative or the only view point of the serviceprovision to the ground controllers.

The utility base 4, or a third satellite can serve as relay points, butthese constitute less preferred options.

When the service vehicle 6 is in close proximity to the targetspacecraft 2 even the telemetry/telecommand link can be performedthrough the target spacecraft 2. The method for receiving telecommandsat the service vehicle 6 in this case is by listening to the telemetryof the target spacecraft 2 and select those packets that will beproperly identified that are addressed to the service vehicle 6. Thiswill further reduce the energy waste and increase the comfort of thetarget spacecraft 2 operators.

Apart from the cases where the service vehicle 6 will act alone or withthe help of a service vehicle 6 a set or behaviors is designed tocapitalize on the fact that a plurality of them will be available.

A method for reaching a signal from a remote place back to the utilitybase 4 or elsewhere can be performed by placing a plurality of servicevehicle 6 in distances according to their respective telecommunicationsmeans and effect the transmission by means of relaying from one to theother the signal until it reaches the destination.

A service vehicle 6 also can carry other service vehicle 6 (towingpushing) docking side by side.

A set of service vehicles 6 can add on their thrust power and perform arelocation mission.

A set of service vehicles 6 can add their reception transmission meansin a formation of a large phased antenna array by positioning themselvesaccording to the desired source of signal or target and coordinated bymeans of a special Escort agent of the utility base 4 to operate on thismode.

Several functions may be automated. Most importantly, the dockingoperation to the utility base 4 and the docking operation to the EngineModule. Advantage of both is the reduced need for teleoperators andresources to establish the video and control link.

In the case of the docking to engine module or other service vehicle 6or specialized vehicle 10 which is far apart from the utility base 4 theadditional advantage is the autonomy achieved. It can be planned at anytime. Low level of resources required as docking is performed withoptimum fuel usage and provides high level of confidence to the ownersof the target spacecraft 2.

A currently preferred embodiment of the service vehicle 6 is a canonical(rectangular, pentagonal, hexagonal) rod shaped structure covered withsolar panels. In another embodiment a pair of solar panels shall bedeployable and retractable. When the panels are retracted and secured onthe service vehicle 6 surface the service vehicle 6 can be navigated asa spin axis stabilized spacecraft. The solar panels will be deployedmainly after docking to a target spacecraft 2 to extend beyond the shadeof the satellite that is serviced. The service vehicle 6 will have themain thruster in its bottom side while at the top side will have thesimple grabble mechanism to grabble the target satellite by the interiorof the fuselage.

The one side of the service vehicle 6 will be capable of performingdocking to the utility base 4 or to an Escort vehicle 10 for refueling.The docking and refueling mechanism will be positioned to lower halfpart of the service vehicle 6 so that the refueling can be possible evenif the service vehicle 6 is attached to a target spacecraft 2.

The service vehicle 6 will be passive as regards the mechanism for therefueling docking but with adequate passive targeting aid (laserretro-reflectors) to ease proximity and semi or fully automated docking.The service vehicle 6 will benefit from the stability of the commondocking place. In this way they will be able to switch most of theirequipment (momentum wheels, communication payloads, thermal subsystemsaving), reducing their wear and increasing their lifetime (form 100% upto 1000%). There will be economy of resources. Fuel consumption reducedto zero, power consumption will be reduced to 2%.

The proximity of the service vehicle 6 s one to the other can reduceheat dissipation. Further economy. The proximity of the service vehicles6 can provide inter-alia protection against debris.

The service vehicles 6 can benefit from a deep-storage mode where someelements could even be extracted for placement under special conditionsfor extending their lifetime. The battery can be stored separately formthe service vehicle 6 in appropriate conditions. The fuels can beflushed out to avoid corrosion of tanks, pipe lines, valves and otherelements form leaks. The tanks could be depressurized to reducemechanical stress from pressure. The service vehicles 6 can benefit fromservice vehicle 6-to-Client interface reconfiguration available at theutility base 4. The service vehicle 6 will be receptive to interfaceconfiguration changes. It will be possible to change the Communicationspayload and the grabble mechanism to customize according to clientcharacteristics. The service vehicle 6 can benefit from service vehicle6 to ground interface reconfiguration service available at the utilitybase 4. The utility base 4 will have the capability to change theconfiguration characteristics of the service vehicle 6 Interface to theutility base 4. The communication payload may be adjusted depending onthe required down link to be used through an Escort-service vehicle 6,through the utility base 4 or through the target spacecraft 2 orotherwise.

The service vehicle 6 can benefit from mission dependentreconfiguration. The optimum reusability and efficiency will depend onthis capability of the utility base 4 to provide this type ofreconfiguration. For each mission the fuel reserves will be adjusted,the communication payload will be reconfigured. Transceivers ofappropriate strength will be installed and other characteristics will beadjusted (momentum, thruster position)

When a given spacecraft is close to another spacecraft it can capturethe telemetry produced by the first said spacecraft by very simple meansas the transmission takes place customarily with a unidirectionalantenna and at power levels sufficient to reach earth.

The telemetry information is transmitted into standardized packets andusually consists of acknowledgments of commands, parameter values fromvarious sources, memory dumps and simple echo messages. A number ofthese telemetry data packets and specifically these whose content can beforced to particular content by telecommands (like echo telemetry, ormemory dumps of certain areas) can be selected to carry command datathat are addressed to another spacecraft in the range of the telemetryof the first spacecraft.

This method invented can be used by any spacecraft that can listen-in tothe telemetry of the first said spacecraft.

The method is proposed to be exploited by the plurality of apparatuseshere invented and intent to offer services to target spacecraft 2.

This method, provides merit form the technical and economic point ofview. The means used for the first satellite to perform the telecommandlink are reused at no extra cost by a plurality of other satellites in amaster-slave configuration.

Additional merit of the invention in the case where the method isapplied to control plurality of servicing satellites is the assuranceprovided to the target spacecraft 2 owner that no dangerous commands maybe sent to the plurality of the servicing vehicles. He will have fullvissibility and control to the operations of the servicing vehicles.

The method is applied by the current invention to make economies in thetelecommand reception means and power consumption and to reinforce theconfidence to the target spacecraft 2 owners that they have full controlof the process. Method of recovering telemetry information from asatellite whose telemetry means transmit at very low power output orbuffering is required or encrypting the telemetry information isrequired.

It is desired in certain circumstances to listen from close distance tothe telemetry information of the target spacecraft 2 either because thetelemetry transmission means can not produce a high power signal, eitherfor power constraint/preservation reasons or because of problems in thetelemetry transmission means.

Additional reasons for listening in can be the need to store thetelemetry for transmission at a later time. This is especially useful tolow earth orbiting satellites that circulate earth and therefore are notall the time in the field of view of a ground station.

Still another reason is the possible need to encrypt the telemetrybefore transmission, need that became apparent after the design phase ofthe target spacecraft 2.

In all the above circumstances it will be beneficial to provide a meansof retransmitting the telemetry of a target spacecraft 2 at anotherfrequency and at higher power or with a delay or in encrypted mode or inany combination of the above.

The proposed method of invention is the delivery of a service vehicle 6equipped with the appropriate listen-in, possible buffering, possibleencryption and retransmission means preferably to an up-link channel ordirectly to the ground.

The choice of way of establishing the feed link depends on theavailability of the said up-link. If the direct link is the choiceappropriate modification of the standard service vehicle 6 shall beperformed before mission starts at the utility base 4. The appropriatemodifications shall include above standard power generation means, powerconditioning means and telemetry retransmission means.

An uncontrollable target spacecraft 2 that tumbles is very difficult anddangerous to capture because it may damage the spacecraft that attemptsapproach for the capture.

A new method is proposed for stabilizing a tumbling spacecraft asfollows:

A pair or service vehicles 6 is equipped with an add-on dual wiredeployment/retracting system (WDRS), secured in their lower part of oneof their sides. Each of the said WDRSs are equipped with a camera or thepair of service vehicle 6 is escorted by an Escort service vehicle 6with camera and telecommunication means. The length of the wire (rolledin the said WDRS) shall be several hundred meters in order to allowoperation of the escort service vehicle 6 without risk of contaminationagainst the target spacecraft 2. The middle of the wire is equipped witha multi anchor apparatus or a net or simply a loop, whatever the casedefines as more appropriate that would capture the SC if comes to itspath.

Formation flying of the pair of the service vehicles 6 in proper angleshall enable the tumbling target spacecraft 2 to be captured. Dependingon the moment of inertia of the target spacecraft 2, the servicevehicles 6 shall perform well timed, directed and weighted thrustsagainst the force the wire will effect as it folds around the tumblingspacecraft. A third service vehicle 6 shall observe closely the wholeoperation. It shall ease the targeting of the wire capture and determinethe risk of damage to the spacecraft after the capture is achieved todirect properly the tumbling attenuation operation.

In some cases, the transportation of a target spacecraft 2 to higherlatitudes, if it has been stacked below the required altitude, or needto go to far longitudes, or need to implement a high inclinationcorrection or for other reasons, requires high acceleration-decelerationmaneuvers.

The said transportation requires stability of the solar panels to avoiddeformation or damaging them, and to avoid destabilizing libration ofthe said solar panels during acceleration-deceleration phases of thesaid transportation mission.

A simple, low material requiring method, is envisaged in order to securethe solar panels from deformation and libration caused by saidacceleration/decelerations of the said transportation mission

A plurality of service vehicles 6 (minimum one, preferably two, morepreferably three, most preferably five) equipped each with a wiredeployment & retracting system in one side and a sidewise gripe on theirfront side and a plurality (zero or more) of Engine Modules is deployed.The said Engine modules secure themselves with the help of the saidplurality of service vehicle 6 to the fuselages of the said targetspacecraft 2. Then, each of the service vehicle 6 in turn secures at theEMs the tip of a wire protruding from the said wiredeployment/retracting system. The said service vehicle 6 capture thesolar arrays from their tips at the two ends in a manner that the axisof the body of the said service vehicle 6 is perpendicular to the panelsurface. After securing the grips the wire retracting systems retractthe wires forcing the tips to stability and pressing the lower part ofthe Engine Module/service vehicle 6 against the said target spacecraft2. In this configuration the service vehicle 6 that are attached to thepanel tips can perform thrusts, of which thrusts the vertical componentvector of force is effected mainly to the base of the Engine Module andpartly to the stiffened solar array panels. Advantageously, thedistribution of the force in the three extreme points of the transportedbody gives excellent moment of inertia and steering capabilities.

Steering of the panels can add to the maneuverability of the system.

The thrust history of all thrusters in the system will be archivedtogether with loads (wet or dry), attitude and gyroscopic information,internal acceleration measurements and acceleration measurements asexternally observed by laser ranging from the utility base 4. Thetotality of this information will be analyzed after every mission andnew calibration parameters will be made available. The same parametersminus the ranging information (when away from the utility base 4) willbe monitored real time by the thruster owning object for updating therelative efficiency thruster table.

For the mass calculation the following method applies when measurementtakes place away from the utility base 4. A service vehicle 6 withrecently calibrated thrusters attaches to the target spacecraft 2. Thesolar panels of target spacecraft 2 are secured in the most stable way.A plurality of EA with cameras and ranging lasers position themselves inthe space in front of the target spacecraft 2 a little above and alittle below its expected trajectory at a distance appropriate for thelaser means. They point the laser beams towards the target spacecraft 2and body and they take measurements during a smooth gradual accelerationphase until a few seconds after stopping acceleration. The accelerationshall be smooth and gradual in order to minimize the sloshing of the drymass.

The analysis of thrusts data, ranging data, visual data, and simulationanalysis on ground can give accurate estimation of the total mass andwet mass specifically.

The deployment of the servicing system 1 is proposed to start with thelaunch of a single service vehicle 6 that will make use of the targetspacecraft 2 as a relay point therefore not needing neither escortservice vehicle 6 for the HBTL nor utility base 4. It may be followed byone or more service vehicle 6 and/or by an escort service vehicle 6 withrefueling capabilities. The refueling escort-service vehicle 6 willprovide the required fuel reserves for the current and part of theupcoming fleet. A possible further refueling escort-service vehicle 6may precede the arrival of the utility base 4.

Advantages of this deployment plan is the low initial cost and the highfinal functionality.

Three deployment areas are foreseen in the beginning

-   -   The Geostationary ring    -   The Low earth orbiting satellites    -   The Medium Earth orbits

The invention is presented to start providing service in thegeostationary ring but the similar apply for the lower to earth orbitsand to further missions around other celestial objects or totrajectories between celestial objects.

This split of functionality between utility base 4, service vehicle 6,EM and EA provides for low mass, low cost, high fuel/dry mass ratio,high maneuverability, long range and operating duration in the servicevehicle 6, EA and EM part. On the other had the utility base 4 gives tothe system high reusability, maintainability, multiple uses, eliminationof waste. The system in total provides for efficient, reliable and lowcost service operations.

Main advantage of this architecture is that the service vehicle 6results in an extremely low dry mass, low cost, agile spacecraft thatcan service target spacecraft 2 which require large delta velocitypotential. Yet main advantage of this element of design is that a dualarm robotic facility is also made available in the context of the system(through the utility base 4 component) allowing for extensive servicingoperations.

A particular advantage of this configuration is that the service vehicle6 is released by the highly demanding subsystem budgets (performancecharacteristics), required for a link with earth, which are requiredonly for a small fraction of the lifetime of the service vehicle 6 whilein the rest of the life time represent dead mass (large overhead inmaneuvers). Placing this functional requirement to another element ofthe system that does not perform demanding maneuvers (to the utilitybase 4) it gives high flexibility and low construction and operationalcosts at the service vehicle 6 part. This fundamental characteristic ofthe design of the service vehicle 6 is new, unique and useful.

The service vehicle 6 does not need to have redundancy of most of itssub-systems (power, solar, propulsion). Its only safety characteristicwill be that it will have fail-safe mechanism of its grabble. Theservice vehicle 6 will capitalize on the presence of utility base 4 inthe relative proximity and also of the similar service vehicle 6 thatwill be able to perform a rescue operation with target the failedservice vehicle 6.

Special Escort-service vehicle 6 will have capability to refuel otherservice vehicles 6.

Advantages are: A service vehicle 6 can perform of a heavy mission (highdelta velocity) without having to return to the Utility base forrefueling and performing again the rendezvous with the servingspacecraft (mostly manual and difficult task). Instead it can remainattached to its mission and wait for successive installments of fuel bya refueling service vehicle 6 (depending on availability). In this waythe required wet mass at the beginning of its mission can be verylimited facilitating the rendezvous and docking as well as reducing thecost of orbit maintenance. In the occasion the mission finally requiredreplenishment of the fuel this is achieved by the special Escort-servicevehicle 6.

If a service vehicle 6 runs out of fuel the Escort-service vehicle 6 canreplenish and then either separate or perform flight attached one to theother reducing the risk in case of failure of one of the two. Thespecial-service vehicle 6 in the beginning of the deployment of thesystem may substitute the utility base 4.

The service vehicle 6 will take advantage of the capabilities of theutility base 4 to perform reconfiguration operations. It will be able tochange communication payload and grabble characteristics in order to fitfor service for a variety of potential target spacecraft 2.

The service vehicle 6 shall be able to enter an idle storage mode whendocked on the utility base 4 or to another service vehicle 6. This willconserve the wear of most subsystems even the structure (by thermalcycles) and reduce the consumption of energy. This will become possibleby the presence f the utility base 4 or an Escort-service vehicle 6.

A simplified version of the service vehicle 6 is the Engine Module thatdoes not have cameras and the like for performing a navigation anddocking. Is put in place on an target spacecraft 2 with the help of aservice vehicle 6 or EA and remains there to perform station keeping andinclination maneuvers until it will require fuel replenishment. In thiscase, a service vehicle 6 with capability of automatic docking on theEngine Module will dock and provide fuel for another term of themission.

REFERENCE NUMERALS

-   1 servicing element-   2 target spacecraft (Utility Agent, UA)-   4 utility base-   6 service vehicle-   8 engine module-   10 specialized vehicle-   12 control module-   14, 16 arrows-   20 main body-   22 propulsion system-   24 docking system-   25 exhaust system-   26 cameras-   28 built-in communication system-   30 control system-   32 dashed line-   34 receiver-   36 emitter-   38 arrow-   40 hollow axle-   42 action axle-   44 fail-safe mechanism-   46 double arrow opening tip-   48 surface-   50 nozzle ring-   52 exhaust channel-   54 combustion chamber-   60 communication module-   62 multiplexer-   64 modulator-   66 transmitter-   68 controller module-   70 encoder-   72 camera-   74 proximity sensor-   76 arrow-   78 control module-   80 receiver-   82 arrow-   84 demodulator-   86 docking subsystem-   88 branch line-   90 retroreflectors-   92 refueling module-   100 main body-   102 equipment/storage bay-   104 delivery/refueling rack-   106 storage module-   108 primary solar panels-   110 secondary solar panels-   112 support grid-   114 robotic arm-   116 reflectors-   118 docking station-   120 actual arm-   122 wrist element-   130 pressurization controlled prothalamus-   132 outer doors-   134 internal doors-   140 three-arm small-sized robots-   142 horizontally secured axis-   144 axis-   146 conductive surfaces-   148 human-like manipulation arms-   150 arm-   152 two finger gripper-   154 table-   156 storage racks-   160 utility outlets-   162 tanks-   166 side surface-   168 upper surface-   170 mechanical interface-   172 bridge-   180 catch system-   182 double reel-unreel mechanism-   184 freedom mechanism-   186 conductive tapes-   188 end piece-   190 gripping wrist-   192 gyroscopic acceleraometer-   194 momentum wheels-   196 power conversion box

1-16. (canceled)
 17. A service vehicle for performing an in-space operation on a selected target spacecraft, comprising: a communication module having at least one of a transmission and a receiving characteristic configurable in order to meet at least one of a transmission and a receiving parameter of the selected targeted spacecraft.
 18. The service vehicle as recited in claim 17, wherein the communication module includes a transmitter.
 19. The service vehicle as recited in claim 17, wherein the communication module includes a configurable receiver.
 20. The service vehicle as recited in claim 19, wherein the receiver has a wording frequency that is adjustable in so as to enable to communication with a telemetry channel of the selected target spacecraft.
 21. The service vehicle as recited in claim 20, further comprising a control module configured to provide a setpoint for an output power of the communication module.
 22. The service vehicle as recited in claim 21, further comprising a position sensor connected to an input portion of the control module, the first postion sensor delivering a set of data indicative of a current position of the service vehicle.
 23. The service vehicle as recited in claim 22, further comprising a second position sensor connected to the input portion of the control module, the second position sensor delivering a set of data indicative of a current position of the target spacecraft.
 24. The service vehicle as recited in claim 21, further comprising an orientation sensor connected to the intput portion of the control module, the orientation sensor delivering a set of data indicative of a current orientation of the target spacecraft relative to the service vehicle.
 25. The service vehicle as recited in claim 17, further comprising a docking system having a hollow first axle and a second axle moveably disposed inside the first axle, the second axle carrying an activateable arrow tip.
 26. The service vehicle as recited in claim 17, further comprising an identification device configured to identifying said target spacecraft.
 27. A servicing system for providing in-space service operations to a selected target spacecraft, comprising: a service vehicle that includes a communication module having at least one of a transmission and a receiving characteristic configurable in order to meet at least one of a transmission and a receiving parameter of the selected targeted spacecraft; a ground control module for delivering operational commands to the service vehicle.
 28. The servicing system as recited in claim 27, wherein the ground control module is configured to receive data from the service vehicle using the target spacecraft as a relay station for signals emitted from the service vehicle.
 29. The servicing system as recited in claim 27, further comprising an orbit-based utility base for said service vehicle.
 30. The servicing system as recited in claim 27, further comprising a relay module for forwarding transmitted signals to the service vehicle.
 31. The servicing system as recited in claim 27, wherein further comprising an engine module attachable to at least one of a service agent, the service vehicle, and the target spacecraft.
 32. A method for in-space servicing of a selected target space-craft, the method comprising: performing an in-space operation on the target spacecraft using a service vehicle having a communication module that includes at least one of a transmission and a receiving characteristic configurable in order to meet at least one of a transmission and a receiving parameter of the selected targeted spacecraft; and relaying command signals to the service vehicle using a telemetry channel disposed between a ground control module and the target spacecraft. 